The present invention relates to cooling gas turbine engine components, and more particularly, to cooling gas turbine engine components at or near exhaust system hot gas flowpaths.
Gas turbine engines generally include an exhaust system located at an aft end of the engine. These exhaust systems can include a turbine exhaust case (TEC) that is located aft of the turbine section or sections of the engine. In low-bypass ratio engines and engines for military applications, the TEC is important for straightening hot gas flow for an afterburner system (located aft of the TEC), for improving the engine's radar profile, etc. TEC assemblies generally include a forward outer diameter ring (FODR), located at a forward portion of the TEC, that defines a portion of a hot gas flowpath.
Probes (or sensors) may be positioned to extend through the FODR during engine testing and during regular flight cycles. These probes extend into the hot gas flowpath in order to gather desired data. Often, such probes are positioned in a hole defined through the FODR, and a boss assembly having a “slider” seal plate is positioned to seal a gap formed between the probe and the edges of the hole in the FODR. The slider seal plate can be retained by tabs located at a radially outer surface of the FODR.
Exhaust system components are often subject to adverse pressure gradients and high temperature levels during operation. Those conditions can lead to undesirable stress, wear and damage to engine components. Over time, this can lead to relatively short lifespans of affected components, and lead to significant expenditures of time, effort and money to repair or replace those affected components. In addition, inadequate pressurization of a FODR plenum at the hot gas flowpath can produce a negative pressure difference, and lead to undesirable inflow of hot gases into the FODR plenum at the slider seal plate of the boss assembly and other locations.